Certain components, in particular in the combustor and turbines, of a gas turbine engine are subject, in operation, to high temperature gas flows. In some cases the high temperature gas flows are at temperatures above the melting point of the component material. In order to protect the components, and in particular the surface of the components adjacent to the high temperature gas flows, from these high temperatures, various cooling arrangements are provided.
Generally such arrangements utilise relatively cool compressed air, which is bled from the compressor section of the gas turbine engine, to cool and protect the components subject to the high operating temperatures.
A well known method of cooling and protecting gas turbine components from the high temperature gas flows is film cooling in which a film of cooling air is provided along the surface of the component exposed to the high temperature gas flows. The film of cooling air is produced by conducting a flow of cooling air through a plurality of passages which perforate the wall of the component. The air exiting the passages is directed, by the passages, to flow in a boundary layer along surface of the component. This cools the wall of the component exposed to the high temperature gas flow and provides a protective film of cool air between the high temperature gas flow and the component surface. The protective film assists in keeping the high temperature gas flow away from the surface of the component wall.
The arrangement and configuration of the passages are carefully designed to provide, and ensure, an adequate boundary layer flow of cooling air along the surface of the component. The passages are accordingly generally angled in the flow direction of the hot gas stream so that the cooling air flows in a downstream direction over the surface of the component.
Ideally it is desired that the boundary layer should flow over substantially the entire surface of the component downstream of the passages. However it has been found that the cooling air leaving the passage exit generally forms a cooling stripe no wider than, or hardly wider than, the dimension of the exit of the passage. Limitations on the number, size, and spacing of the passages results in gaps in the protective cooling layer provided and/or areas of reduced protection/cooling.
To overcome this it has been proposed, in for example U.S. Pat. No. 3,527,543, to use divergent passages where the cross section of the passages increases towards the passage exit at the surface of the component exposed to the hot gas flow. The cooling air which flows through the passages is thereby partially spread out over a larger area of the surface. Whilst this is an improvement over a constant cross section passage it has been found that the air exiting the passage generally still does not spread out enough to provide a continuous film of cooling air between the typical spacing of the passages.
A further development of the diverging passages is to arrange the passages sufficiently close to each other such that the outlets of the adjacent passages, on the surface of the component exposed to the hot gas flows, intersect laterally to define a common outlet in the form of a laterally extending slot. The cooling air expands as it passes though the passages and exits from this common slot as a substantially continuous film. Such an arrangement is described more fully in U.S. Pat. No. 4,676,719 which also references other similar arrangements which are described in U.S. Pat. No. 3,515,499 and Japanese Patent Number 55-114806.
In these prior art arrangements the passages are divergent and the cross sectional area of the passage increases towards the exit. This slows down, and diffuses, the flow of cooling air therethrough. As is taught in the prior art this slowing of the flow is important in assisting in spreading the flow of cooling air, in a boundary layer, along and over the surface of the component. Another important consideration in the design of such film cooling arrangements is to ensure that a stable boundary layer is provided over the surface of the component, and that this boundary layer remains attached to the surface of the component to thereby protect the surface from the high temperature gas stream. This boundary layer flow of cooling air is also required to withstand fluctuations and variations in the hot gas stream, that may occur during operation, to ensure that adequate cooling and protection is provided throughout the operation of the engine. In addition the flow through the passages and along the surface of the component should be as aerodynamically efficient as possible.
In an additional variation slots within the walls of the component can be used to direct the cooling air to the outer surface of the component. Such an arrangement is described in U.S. Pat. Nos. 2,149,510, 2,220,420 and 2,489,683.
Although such arrangements provide a good flow of cooling air along and over the surface of the component the structural strength of the walls of the component is reduced. This is also true, albeit to a lesser extent, with the arrangements where the passages intersect at their exits to form a common exit slot.
It is therefore desirable to provide an improved gas turbine engine component cooling arrangement and configuration, and in particular to provide an improved arrangement and configuration of cooling passages that address the above mentioned problems and/or offers improvements to such cooling arrangements generally.